WO2005106208A1 - Pale de turbine a gaz - Google Patents

Pale de turbine a gaz Download PDF

Info

Publication number
WO2005106208A1
WO2005106208A1 PCT/EP2005/051721 EP2005051721W WO2005106208A1 WO 2005106208 A1 WO2005106208 A1 WO 2005106208A1 EP 2005051721 W EP2005051721 W EP 2005051721W WO 2005106208 A1 WO2005106208 A1 WO 2005106208A1
Authority
WO
WIPO (PCT)
Prior art keywords
cooling
blade
shroud
region
bores
Prior art date
Application number
PCT/EP2005/051721
Other languages
German (de)
English (en)
Inventor
Ulrich Rathmann
Original Assignee
Alstom Technology Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Alstom Technology Ltd filed Critical Alstom Technology Ltd
Priority to CN2005800138966A priority Critical patent/CN1950589B/zh
Priority to AT05747380T priority patent/ATE551497T1/de
Priority to AU2005238655A priority patent/AU2005238655C1/en
Priority to EP05747380A priority patent/EP1740797B1/fr
Publication of WO2005106208A1 publication Critical patent/WO2005106208A1/fr
Priority to US11/549,767 priority patent/US7273347B2/en

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/10—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the invention relates to a blade for a gas turbine and in particular to cooling for the shroud of the blade.
  • Shrouds for gas turbine blades are used to seal and limit the leakage flow in the gap area between the blade tips and the radially opposite stator or rotor. They extend in the circumferential direction and over a certain area in the direction of the turbine axis, if possible in adaptation of the contour of the inner housing or of the rotor.
  • Conventional shrouds in many cases also have one or more sealing ribs, also called cutting edges, to improve the seal, which extend from a platform of the shroud, i.e. a substantially flat section of the shroud, along the radial direction.
  • the cover tapes are convectively cooled in order to extend their operating time in the gas turbine through which hot gas flows, as disclosed for example in EP 1013884 and EP 1083299.
  • a blade with a shroud is described there, which has several bores for a cooling air flow.
  • Bores are connected to a cooling channel in the airfoil and each lead to a lateral outlet in the circumferential direction.
  • EP 1041247 discloses a gas turbine blade with inner, radially extending cooling channels which open into a plenum 42 and 44. From there, bores 54, 56, 58 extend in the plane of the shroud, through which the shroud is cooled by means of film cooling and convective cooling. In a variant, the bores extend obliquely from the plenum and in a slightly radial direction to the radially outer surface of the shroud platform.
  • a cover band of a gas turbine blade is subjected to different thermal loads along the direction of flow of the hot gas and is also subjected to different mechanical loads in different areas.
  • the requirements regarding cooling and mechanical resilience are also in different areas of the shroud different. This is taken into account in the aforementioned gas turbine blades by adapting the bore diameter and other measures to change the pressure differentials.
  • the cover band of a gas turbine blade extends in the circumferential direction along the blade tip and in the radial direction with respect to the turbine rotor and is arranged opposite a stator housing.
  • the cover band is divided into areas that are subjected to different thermal loads.
  • the different areas are cooled by different cooling arrangements, each cooling arrangement allowing cooling with a different physical effect that is adapted to the thermal load, such as film cooling, impingement cooling, convective cooling or mixed cooling.
  • the gas turbine blade has a first cooling arrangement for cooling a first region of the shroud by cooling air from a cooling system from the interior of the blade.
  • This first area is the first area in the direction of the hot gas flow and is therefore most thermally stressed.
  • a second area downstream from the first area in the direction of the hot gas flow is thermally less stressed than the first area.
  • Radially opposite one of the gas turbine blades arranged stator the second cooling arrangement is arranged, which serves to cool the second area of the shroud from outside the blade.
  • the first and second cooling arrangements are different from one another in that the first cooling arrangement effects a covenant and film cooling and the second cooling arrangement effects an impingement cooling.
  • the cooling of the cover band according to the invention brings about a cooling which is appropriate for the thermal load of the areas and a correspondingly appropriate cooling air consumption.
  • the first region of the cover band of the gas turbine blade has, in particular, a cutting edge which extends in the radial direction with respect to the gas turbine rotor and runs in the longitudinal direction in the circumferential direction and in which the first cooling arrangement is arranged.
  • the cutting edge has a plurality of bores which are in flow connection with a cooling channel of the airfoil and have outlets on the hot gas side of the shroud.
  • a flow of cooling air causes convective cooling of the cutting edge as it flows through the bores. After exiting the holes, it flows along the outer surface of the shroud and causes film cooling there.
  • the stator housing which is arranged radially opposite the shroud, has a plurality of cooling channels which are directed essentially perpendicular to the platform of the shroud. They are used to cool the second area of the third Cover bands in the hot gas flow direction. They are connected to the stator cooling system, as a result of which cooling air, which is branched off, flows through the cooling ducts onto the platform of the cover band and effects impact cooling there. The cooling air then escapes in both axial directions, whereby a blocking flow in the opposite direction to the leakage flow can occur.
  • the second area of the shroud is delimited on both sides in the axial direction by radially running cutting edges.
  • the gas turbine blade has a further third region of the shroud in the direction of the hot gas flow, which is equipped with a third cooling arrangement.
  • This cooling arrangement has a plurality of bores which are in flow connection with a cooling channel in the interior of the airfoil. The bores are directed at an angle to the radial in at least a partial radial outward direction that direct a cooling air flow to the radially outer portion of the shroud. Cooling air flowing through these holes causes convective cooling of this third area.
  • the bores in the plane of the shroud platform are oriented at an angle with respect to the circumferential direction so that the cooling air is blown out of the bores essentially counter to the direction of rotation of the blades.
  • the bores in the end area run parallel to one another.
  • a plurality of further cooling channels are arranged in the stator radially opposite the shroud, which are directed essentially perpendicularly to a third region of the shroud in the direction of the hot gas flow. They are used to cool this third area.
  • the third area is delimited by a cutting edge in the axial direction and in the opposite direction of the hot gas flow.
  • the cooling ducts are in flow communication with the cooling system of the stator, as a result of which cooling air from the stator cooling system is directed onto the end region of the cover band and effects impingement there.
  • FIG. 1 shows a section through a rotating gas turbine blade and part of the opposite stator with a cooling arrangement according to the first and second embodiment of the invention
  • FIG. 2 shows a top view of the cover band of the gas turbine blade
  • Figure 3 is a side view of the shroud along the line HI-HI, for
  • Figure 4 is a view of the shroud along the section according to IV-IV
  • FIG. 5 is a detailed view according to V in Figure 4 to show a preferred
  • Figure 6 shows a section through a rotating gas turbine blade as in Figure 1 with a cooling arrangement according to the third embodiment of the invention.
  • FIG. 1 shows a rotating gas turbine blade in a meridional section through the gas turbine.
  • the directions x and z indicate the axial direction, that is to say the direction of the machine axis, or the radial direction with respect to the gas turbine rotor.
  • the airfoil 1 is shown and the airfoil tip on which the shroud 2 is arranged.
  • the stator housing 4 is shown opposite the shroud 2, in the radial outward direction with respect to the gas turbine rotor 3.
  • the gas turbine blade and the stator housing each have a cooling system 5 and 6, respectively.
  • the direction of the hot gas flow is marked with an arrow 7. Basically, the temperature of the hot gas flow and correspondingly the thermal load on the machine components along direction 7 decreases continuously.
  • the cover tape 2 is divided into three areas A, B and C.
  • the first area A is exposed to a higher temperature of the hot gas flow in comparison to the two subsequent areas B and C and is consequently most thermally stressed.
  • the first area has a cutting edge 8 which extends radially outwards and in the circumferential direction.
  • the cutting edge 8 has a bore 9 which is in flow connection with the
  • Cooling system 5 is. This bore extends, for example, in the circumferential direction within the cutting edge. Several further bores 10 branch off from this bore 9 and extend radially inwards until they emerge on the rotor-side surface of the cutting edge, that is to say on the hot gas side of the shroud. The branching bores 10 are shown in FIG. 3. Cooling air from the cooling system 5 of the airfoil flows through the bore 9 and through the branching bores 10, causing convective cooling of the cutting edge 8. The exits of the bores are each designed in such a way that cooling air escapes along the surface of the cutting edge and causes additional film cooling there. The cutting edge is thus cooled by two different cooling mechanisms.
  • a cooling channel 11 is arranged through the wall of the housing 4, which is connected to the cooling system in the stator housing.
  • a cooling air flow indicated by the arrow 12 flows from this cooling system through the cooling channel 11 and, due to its orientation, is preferably directed perpendicularly onto the shroud 2.
  • the Cooling channel 11 also aligned at a different angle with respect to the shroud.
  • the cooling air flow 12 thus effects an impact cooling of the central region B of the shroud.
  • the area B is delimited in the axial direction and in the direction of the hot gas flow by the first cutting edge 8 and a second cutting edge 13.
  • the cooling air flow 12 escapes from the limited area as a leakage flow in that the cooling air flow flows away in both axial directions via the cutting edge 8 and the cutting edge 13. Depending on the operating conditions, a blocking flow against a hot gas leakage flow can result. Usually, due to degradation effects, mixed cooling of the shroud will result over time.
  • a special opening or gap is provided in the area of the second sealing cutting edge 13, which enables the cooling air to flow out in a precisely controlled manner.
  • a plurality of bores are arranged which originate from the cooling system 5 of the airfoil and run to the radially outer surface of the shroud. A flow of cooling air through these holes causes convective cooling of this area. They are shown in Figure 2.
  • FIG. 2 shows a plan view of the cover band according to the invention, again with the areas A, B and C.
  • the axial direction and the circumferential direction with respect to the turbine rotor are shown with x and y, as well as the outline of the blade root 14 and the outline of the blade itself with a broken line
  • the cutting edge 8 in area A and the cutting edge 13 in area B are shown, which run in the circumferential direction and serve to seal against leakage currents.
  • the area C has the bores 15 for the purpose of convective cooling of that area, wherein they run at an angle a to the circumferential direction y.
  • the angle ⁇ is, for example, in a range between 2 ° and 90 °.
  • the cooling air that emerges from the bores 15 is blown out in the opposite direction to the direction of rotation of the blade.
  • the bores 15 are preferably aligned parallel to one another, so that production is simplified.
  • FIG. 3 shows a section according to HI-HI in FIG. 2 and shows the cutting edge 8 in area A of the shroud and the course of the transverse bore 9 and the bores 10 branching from it.
  • the transverse bore 9 is in flow connection with the cooling system of the airfoil via channel 21 connected.
  • the Flow connection is ensured by an expansion of the cooling system of the airfoil, which extends into the cutting edge 8 and opens into the transverse bore 9.
  • the plurality of branching bores 10 run substantially radially inward with respect to the turbine rotor to exit on the hot gas side of the cutting edge 8.
  • the course of the cooling flow is indicated by arrows through the channel 21, via the transverse bore 9 and the branching bores 10.
  • the exits from the bore 10 are designed, in particular, to effect film cooling of the hot gas-side surface of the cutting edge, for example with a slightly diverging exit part and a preferred angular range, as is known from the relevant literature.
  • Preferred methods of production are the usual casting processes with a core and drilling from the outside and subsequent closing of the borehole entrances by means of plugs 20, which e.g. be inserted in a form-fitting manner or connected materially (soldering, welding).
  • FIG. 4 shows the configuration of the bores 15 in a section according to IV-IV.
  • the blade and a channel of the cooling system 5 are shown in its blade.
  • the bore 15 extends from the channel and extends to the radially outer surface of the shroud 2.
  • the exit of a bore 15 is designed to be angled so that the mixture with the hot gas flow can be advantageously influenced according to the conditions.
  • the angle ⁇ between the exit surface and the axis of the bore is preferably in a range between 40 ° and 140 °.
  • the angle ⁇ between the exit surface and the direction of the radial z is preferably selected in a range from 30 ° to 120 °.
  • the diameter of the bore is in a range between 0.6 and 4.5 mm, preferably in a range between 0.6 and 2.5 mm. This is for adequate convective cooling for this area.
  • FIG. 5 shows in a section according to IV-IV a variant of the exit of the bores 15.
  • the exit surface is again angled and stepped with respect to the bore axis, the end of the upper lip 16 being essentially perpendicular to the bore axis.
  • the dimension s depends on the diameter of the exit surface and is in particular in a ratio of 0.5 to 3 in relation to the diameter of the bore and also allows the mixture with the hot gas flow to be advantageously influenced.
  • FIG. 6 shows, in the same meridional section as in FIG. 1, a gas turbine blade 1 according to the third embodiment of the invention.
  • an additional channel is arranged in the stator housing, through which cooling air from the cooling system of the housing is directed onto the shroud. Impact cooling is effected there as for area B.
  • the gas turbine blade is coated completely or in individual areas with a heat barrier layer in accordance with its use in the gas turbine.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne une pale (1) d'une turbine à gaz présentant une sangle (3) qui est refroidie dans différentes zones (A,B,C) en fonction de la charge thermique et par différents mécanismes de refroidissement. Dans une première zone (A), une lame (8) présente des orifices qui permettent un refroidissement par convection de la lame et un refroidissement par film du côté gaz chaud de la lame. Une deuxième zone (B) est refroidie par impact par un flux d'air froid provenant d'un canal du boîtier stator opposé. Une troisième zone (C) présente plusieurs orifices de disposition parallèle qui s'étendent d'un canal de refroidissement d'un système de refroidissement destiné de l'ailette jusqu'à la surface externe radiale de la sangle. Un flux d'air de refroidissement qui s'écoule par ces orifices entraîne un refroidissement par convection de cette zone.
PCT/EP2005/051721 2004-04-30 2005-04-19 Pale de turbine a gaz WO2005106208A1 (fr)

Priority Applications (5)

Application Number Priority Date Filing Date Title
CN2005800138966A CN1950589B (zh) 2004-04-30 2005-04-19 燃气轮机的叶片
AT05747380T ATE551497T1 (de) 2004-04-30 2005-04-19 Gasturbine
AU2005238655A AU2005238655C1 (en) 2004-04-30 2005-04-19 Blade for a gas turbine
EP05747380A EP1740797B1 (fr) 2004-04-30 2005-04-19 Turbine a gaz
US11/549,767 US7273347B2 (en) 2004-04-30 2006-10-16 Blade for a gas turbine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP04101876.3 2004-04-30
EP04101876A EP1591626A1 (fr) 2004-04-30 2004-04-30 Aube de turbine à gaz

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US11/549,767 Continuation US7273347B2 (en) 2004-04-30 2006-10-16 Blade for a gas turbine

Publications (1)

Publication Number Publication Date
WO2005106208A1 true WO2005106208A1 (fr) 2005-11-10

Family

ID=34929047

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2005/051721 WO2005106208A1 (fr) 2004-04-30 2005-04-19 Pale de turbine a gaz

Country Status (8)

Country Link
US (1) US7273347B2 (fr)
EP (2) EP1591626A1 (fr)
KR (1) KR20070006875A (fr)
CN (1) CN1950589B (fr)
AT (1) ATE551497T1 (fr)
AU (1) AU2005238655C1 (fr)
MY (1) MY142730A (fr)
WO (1) WO2005106208A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7273347B2 (en) 2004-04-30 2007-09-25 Alstom Technology Ltd. Blade for a gas turbine

Families Citing this family (32)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2003291B1 (fr) * 2007-06-15 2017-08-09 Ansaldo Energia Switzerland AG Aube de turbine à gaz moulée et procédé de production
EP2009248B1 (fr) * 2007-06-25 2010-05-12 Siemens Aktiengesellschaft Agencement de turbine et procédé de refroidissement d'un anneau situé au bout d'une aube de turbine
KR100857346B1 (ko) * 2007-07-06 2008-09-05 현대자동차주식회사 Ｓｏｏｔ 영향을 고려한 Ｕｒｅａ 분사 방법
US8182207B2 (en) * 2008-03-17 2012-05-22 General Electric Company Inner turbine shell support configuration and methods
CH699593A1 (de) 2008-09-25 2010-03-31 Alstom Technology Ltd Schaufel für eine gasturbine.
DE102009049649A1 (de) * 2009-10-15 2011-04-21 Abb Turbo Systems Ag Turbinenrad
US8721289B2 (en) * 2009-10-30 2014-05-13 General Electric Company Flow balancing slot
US8579581B2 (en) * 2010-09-15 2013-11-12 General Electric Company Abradable bucket shroud
RU2547542C2 (ru) * 2010-11-29 2015-04-10 Альстом Текнолоджи Лтд Осевая газовая турбина
RU2547541C2 (ru) * 2010-11-29 2015-04-10 Альстом Текнолоджи Лтд Осевая газовая турбина
JP2012154201A (ja) * 2011-01-24 2012-08-16 Ihi Corp タービン動翼及びシール構造
US20120195742A1 (en) * 2011-01-28 2012-08-02 Jain Sanjeev Kumar Turbine bucket for use in gas turbine engines and methods for fabricating the same
US9249670B2 (en) 2011-12-15 2016-02-02 General Electric Company Components with microchannel cooling
US9109455B2 (en) * 2012-01-20 2015-08-18 General Electric Company Turbomachine blade tip shroud
US20130318996A1 (en) * 2012-06-01 2013-12-05 General Electric Company Cooling assembly for a bucket of a turbine system and method of cooling
EP2713009B1 (fr) * 2012-09-26 2015-03-11 Alstom Technology Ltd Procédé et système de refroidissement pour refroidir des aubes d'au moins une rangée d'aubes dans une turbomachine rotative
GB201308605D0 (en) 2013-05-14 2013-06-19 Rolls Royce Plc A shroud arrangement for a gas turbine engine
GB201309769D0 (en) * 2013-05-31 2013-07-17 Cummins Ltd A seal assembly
EP2837769B1 (fr) * 2013-08-13 2016-06-29 Alstom Technology Ltd Arbre de rotor pour turbomachine
US9759070B2 (en) * 2013-08-28 2017-09-12 General Electric Company Turbine bucket tip shroud
RU2568763C2 (ru) * 2014-01-30 2015-11-20 Альстом Текнолоджи Лтд Компонент газовой турбины
EP3034790B1 (fr) 2014-12-16 2020-06-24 Ansaldo Energia Switzerland AG Aube rotative pour une turbine à gaz
US10208602B2 (en) * 2015-04-27 2019-02-19 United Technologies Corporation Asymmetric diffuser opening for film cooling holes
US20180179901A1 (en) * 2015-07-24 2018-06-28 Siemens Aktiengesellschaft Turbine blade with contoured tip shroud
US10508554B2 (en) 2015-10-27 2019-12-17 General Electric Company Turbine bucket having outlet path in shroud
US10184342B2 (en) * 2016-04-14 2019-01-22 General Electric Company System for cooling seal rails of tip shroud of turbine blade
US10344599B2 (en) * 2016-05-24 2019-07-09 General Electric Company Cooling passage for gas turbine rotor blade
US10502069B2 (en) * 2017-06-07 2019-12-10 General Electric Company Turbomachine rotor blade
US11060407B2 (en) 2017-06-22 2021-07-13 General Electric Company Turbomachine rotor blade
US10590777B2 (en) * 2017-06-30 2020-03-17 General Electric Company Turbomachine rotor blade
US10301943B2 (en) 2017-06-30 2019-05-28 General Electric Company Turbomachine rotor blade
US10577945B2 (en) 2017-06-30 2020-03-03 General Electric Company Turbomachine rotor blade

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1163559A (fr) * 1956-12-21 1958-09-29 Bertin & Cie Perfectionnement aux turbines
JPS5847104A (ja) * 1981-09-11 1983-03-18 Agency Of Ind Science & Technol ガスタ−ビンのタ−ビン動翼
US5460486A (en) * 1992-11-19 1995-10-24 Bmw Rolls-Royce Gmbh Gas turbine blade having improved thermal stress cooling ducts
DE10336863A1 (de) * 2002-09-17 2004-03-25 Alstom (Switzerland) Ltd. Thermische Turbomaschine

Family Cites Families (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3606574A (en) * 1969-10-23 1971-09-20 Gen Electric Cooled shrouded turbine blade
US4311431A (en) * 1978-11-08 1982-01-19 Teledyne Industries, Inc. Turbine engine with shroud cooling means
US6042878A (en) * 1996-12-31 2000-03-28 General Electric Company Method for depositing a ceramic coating
EP1013884B1 (fr) 1998-12-24 2005-07-27 ALSTOM Technology Ltd Aube de turbine avec plateforme refroidie
EP1041247B1 (fr) 1999-04-01 2012-08-01 General Electric Company Aube de turbineà gaz comprenant un circuit de refroidissement ouvert
US6254345B1 (en) * 1999-09-07 2001-07-03 General Electric Company Internally cooled blade tip shroud
DE10064265A1 (de) * 2000-12-22 2002-07-04 Alstom Switzerland Ltd Vorrichtung und Verfahren zur Kühlung einer Plattform einer Turbinenschaufel
RU2271454C2 (ru) * 2000-12-28 2006-03-10 Альстом Текнолоджи Лтд Устройство площадок в прямоточной осевой газовой турбине с улучшенным охлаждением участков стенки и способ уменьшения потерь через зазоры
AU2002366846A1 (en) * 2001-12-13 2003-07-09 Alstom Technology Ltd Hot gas path subassembly of a gas turbine
EP1591626A1 (fr) 2004-04-30 2005-11-02 Alstom Technology Ltd Aube de turbine à gaz

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1163559A (fr) * 1956-12-21 1958-09-29 Bertin & Cie Perfectionnement aux turbines
JPS5847104A (ja) * 1981-09-11 1983-03-18 Agency Of Ind Science & Technol ガスタ−ビンのタ−ビン動翼
US5460486A (en) * 1992-11-19 1995-10-24 Bmw Rolls-Royce Gmbh Gas turbine blade having improved thermal stress cooling ducts
DE10336863A1 (de) * 2002-09-17 2004-03-25 Alstom (Switzerland) Ltd. Thermische Turbomaschine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
PATENT ABSTRACTS OF JAPAN vol. 0071, no. 28 (M - 220) 3 June 1983 (1983-06-03) *

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7273347B2 (en) 2004-04-30 2007-09-25 Alstom Technology Ltd. Blade for a gas turbine

Also Published As

Publication number Publication date
AU2005238655B2 (en) 2010-08-26
AU2005238655C1 (en) 2011-06-09
EP1591626A1 (fr) 2005-11-02
MY142730A (en) 2010-12-31
ATE551497T1 (de) 2012-04-15
CN1950589B (zh) 2012-02-22
EP1740797A1 (fr) 2007-01-10
CN1950589A (zh) 2007-04-18
US20070071593A1 (en) 2007-03-29
US7273347B2 (en) 2007-09-25
AU2005238655A1 (en) 2005-11-10
EP1740797B1 (fr) 2012-03-28
KR20070006875A (ko) 2007-01-11

Similar Documents

Publication Publication Date Title
EP1740797B1 (fr) Turbine a gaz
EP1320661B1 (fr) Aube de turbine a gaz
EP2123860B1 (fr) Redresseur de tourbillons combiné
EP1267041B1 (fr) Aube de turbine refroidie
EP2414640B1 (fr) Aube pour une turbine à gaz
DE19810821A1 (de) Gasturbinen-Dichtungsvorrichtung
WO2003052240A2 (fr) Systeme de turbine a gaz
EP1260678A1 (fr) Dispositif de refroidissement pour les éléments d'une turbine à gaz
DE102006004437A1 (de) Plattform einer Laufschaufel einer Gasturbine, Verfahren zur Herstellung einer Laufschaufel, Dichtungsplatte und Gasturbine
EP3121373B1 (fr) Roue de turbine refroidie, plus particulièrement pour un réacteur
DE19904229A1 (de) Gekühlte Turbinenschaufel
WO2011029420A1 (fr) Dispositif de déflexion pour un écoulement de fuite dans une turbine à gaz et turbine à gaz correspondante
DE19617539B4 (de) Rotor für eine thermische Turbomaschine
EP3336313A1 (fr) Ensemble d'aube mobile pour turbines d'une turbine turbine à gaz et procédé de fourniture d'air sceau dans un ensemble d'aube mobile pour turbines
EP2823152A1 (fr) Aube mobile de turbine et section axiale de rotor pour une turbine à gaz
EP0992656B1 (fr) Turbomachine pour comprimer ou détendre un fluide comprimable
EP3473808B1 (fr) Pale d'aube pour une aube mobile de turbine à refroidissement intérieur ainsi que procédé de fabrication d'une telle pale
EP1621735B1 (fr) Rotor d'une turbomachine
DE69825964T2 (de) Deckband für gasturbinenschaufelspitzen
EP1288435B1 (fr) Aube de turbine avec au moins un orifice de refroidissement
WO2008155248A1 (fr) Refroidissement de l'aube directrice d'une turbine à gaz
EP3112593A1 (fr) Aube de turbine a refroidissement interieur
DE19940556B4 (de) Vorrichtung zum Kühlen von Leit- oder Laufschaufeln in einer Gasturbine
EP1456507B1 (fr) Agencement d'etancheite pour composants d'une turbomachine
EP2453109B1 (fr) Agencement de turbine à gaz et procédé de fonctionnement d'un agencement de turbine à gaz

Legal Events

Date Code Title Description
AK Designated states

Kind code of ref document: A1

Designated state(s): AE AG AL AM AT AU AZ BA BB BG BR BW BY BZ CA CH CN CO CR CU CZ DE DK DM DZ EC EE EG ES FI GB GD GE GH GM HR HU ID IL IN IS JP KE KG KM KP KR KZ LC LK LR LS LT LU LV MA MD MG MK MN MW MX MZ NA NI NO NZ OM PG PH PL PT RO RU SC SD SE SG SK SL SM SY TJ TM TN TR TT TZ UA UG US UZ VC VN YU ZA ZM ZW

AL Designated countries for regional patents

Kind code of ref document: A1

Designated state(s): GM KE LS MW MZ NA SD SL SZ TZ UG ZM ZW AM AZ BY KG KZ MD RU TJ TM AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IS IT LT LU MC NL PL PT RO SE SI SK TR BF BJ CF CG CI CM GA GN GQ GW ML MR NE SN TD TG

121 Ep: the epo has been informed by wipo that ep was designated in this application
WWE Wipo information: entry into national phase

Ref document number: 2005747380

Country of ref document: EP

WWE Wipo information: entry into national phase

Ref document number: 2005238655

Country of ref document: AU

WWE Wipo information: entry into national phase

Ref document number: 11549767

Country of ref document: US

ENP Entry into the national phase

Ref document number: 2005238655

Country of ref document: AU

Date of ref document: 20050419

Kind code of ref document: A

WWP Wipo information: published in national office

Ref document number: 2005238655

Country of ref document: AU

WWE Wipo information: entry into national phase

Ref document number: 1020067022554

Country of ref document: KR

WWE Wipo information: entry into national phase

Ref document number: 200580013896.6

Country of ref document: CN

NENP Non-entry into the national phase

Ref country code: DE

WWW Wipo information: withdrawn in national office

Country of ref document: DE

WWP Wipo information: published in national office

Ref document number: 2005747380

Country of ref document: EP

WWP Wipo information: published in national office

Ref document number: 1020067022554

Country of ref document: KR

WWP Wipo information: published in national office

Ref document number: 11549767

Country of ref document: US